Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

WORTMANN FX M2 AIRFOIL (fxm2-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: WORTMANN FX M2 AIRFOIL (fxm2-il)
Reynolds number: 100,000
Max Cl/Cd: 45.93 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fxm2-il-100000.txt
Download as CSV file: xf-fxm2-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: WORTMANN FX M2 AIRFOIL                          
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.2698   0.09562   0.09102  -0.0153   1.0000   0.0488
  -7.000  -0.2705   0.09360   0.08908  -0.0146   1.0000   0.0498
  -6.750  -0.2739   0.09175   0.08732  -0.0136   1.0000   0.0509
  -6.500  -0.2752   0.08975   0.08541  -0.0138   1.0000   0.0519
  -6.250  -0.2747   0.08786   0.08361  -0.0159   1.0000   0.0530
  -6.000  -0.2693   0.08628   0.08207  -0.0231   1.0000   0.0538
  -5.750  -0.2617   0.08307   0.07891  -0.0262   1.0000   0.0543
  -5.500  -0.2615   0.07946   0.07540  -0.0201   1.0000   0.0552
  -5.250  -0.2572   0.07676   0.07278  -0.0182   1.0000   0.0565
  -5.000  -0.2506   0.07420   0.07028  -0.0184   1.0000   0.0583
  -4.750  -0.2403   0.07154   0.06764  -0.0206   1.0000   0.0613
  -4.500  -0.2091   0.06820   0.06414  -0.0322   1.0000   0.0648
  -4.250  -0.2124   0.06535   0.06146  -0.0269   1.0000   0.0664
  -4.000  -0.2040   0.06298   0.05916  -0.0263   1.0000   0.0693
  -3.750  -0.1602   0.06002   0.05585  -0.0380   1.0000   0.0757
  -3.500  -0.1462   0.05607   0.05212  -0.0370   0.9959   0.0780
  -3.250  -0.0805   0.05147   0.04724  -0.0494   0.9865   0.0881
  -3.000  -0.0239   0.04752   0.04307  -0.0582   0.9767   0.1004
  -2.750   0.0278   0.04349   0.03898  -0.0652   0.9669   0.1154
  -2.500   0.0722   0.04005   0.03558  -0.0700   0.9530   0.1359
  -1.750   0.2077   0.03120   0.02675  -0.0827   0.9039   0.2923
  -1.250   0.2613   0.02654   0.02259  -0.0778   0.8529   0.4685
  -1.000   0.2932   0.02455   0.02062  -0.0759   0.8167   0.5434
  -0.750   0.4450   0.02086   0.01378  -0.1004   0.7625   0.1390
  -0.500   0.4825   0.01994   0.01218  -0.1002   0.7081   0.1239
  -0.250   0.5139   0.01955   0.01110  -0.0990   0.6583   0.1121
   0.000   0.5411   0.01884   0.01004  -0.0980   0.6134   0.1080
   0.250   0.5675   0.01852   0.00936  -0.0968   0.5733   0.1060
   0.500   0.5933   0.01839   0.00889  -0.0956   0.5390   0.1090
   0.750   0.6187   0.01824   0.00853  -0.0945   0.5057   0.1097
   1.000   0.6436   0.01815   0.00823  -0.0933   0.4785   0.1109
   1.250   0.6690   0.01821   0.00807  -0.0924   0.4543   0.1140
   1.500   0.6950   0.01819   0.00794  -0.0917   0.4322   0.1233
   1.750   0.7222   0.01835   0.00785  -0.0912   0.4154   0.1330
   2.000   0.7500   0.01855   0.00792  -0.0908   0.3999   0.1621
   2.250   0.7667   0.01711   0.00816  -0.0879   0.3872   1.0000
   2.500   0.7937   0.01757   0.00831  -0.0873   0.3743   1.0000
   2.750   0.8206   0.01805   0.00853  -0.0869   0.3633   1.0000
   3.000   0.8475   0.01859   0.00880  -0.0865   0.3538   1.0000
   3.250   0.8743   0.01917   0.00917  -0.0861   0.3441   1.0000
   3.500   0.9007   0.01976   0.00969  -0.0858   0.3344   1.0000
   3.750   0.9274   0.02037   0.01016  -0.0855   0.3265   1.0000
   4.000   0.9537   0.02089   0.01063  -0.0851   0.3185   1.0000
   4.250   0.9799   0.02143   0.01113  -0.0847   0.3111   1.0000
   4.500   1.0064   0.02191   0.01148  -0.0843   0.3046   1.0000
   4.750   1.0320   0.02248   0.01205  -0.0839   0.2974   1.0000
   5.000   1.0589   0.02307   0.01244  -0.0837   0.2917   1.0000
   5.250   1.0839   0.02380   0.01327  -0.0832   0.2852   1.0000
   5.500   1.1106   0.02455   0.01392  -0.0830   0.2800   1.0000
   5.750   1.1348   0.02536   0.01488  -0.0825   0.2740   1.0000
   6.000   1.1606   0.02607   0.01552  -0.0821   0.2681   1.0000
   6.250   1.1838   0.02680   0.01643  -0.0814   0.2613   1.0000
   6.500   1.2098   0.02740   0.01693  -0.0811   0.2559   1.0000
   6.750   1.2322   0.02814   0.01789  -0.0803   0.2491   1.0000
   7.000   1.2579   0.02870   0.01832  -0.0800   0.2437   1.0000
   7.250   1.2804   0.02947   0.01922  -0.0792   0.2369   1.0000
   7.500   1.3046   0.03030   0.02004  -0.0788   0.2311   1.0000
   7.750   1.3272   0.03120   0.02100  -0.0781   0.2245   1.0000
   8.000   1.3485   0.03236   0.02229  -0.0774   0.2182   1.0000
   8.250   1.3707   0.03331   0.02330  -0.0767   0.2121   1.0000
   8.500   1.3898   0.03460   0.02480  -0.0757   0.2065   1.0000
   8.750   1.4101   0.03567   0.02605  -0.0748   0.2011   1.0000
   9.000   1.4330   0.03668   0.02705  -0.0743   0.1971   1.0000
   9.250   1.4469   0.03835   0.02911  -0.0728   0.1919   1.0000
   9.500   1.4680   0.03931   0.03012  -0.0721   0.1882   1.0000
   9.750   1.4937   0.04019   0.03091  -0.0719   0.1854   1.0000
  10.000   1.4986   0.04297   0.03426  -0.0698   0.1818   1.0000
  10.250   1.5083   0.04531   0.03692  -0.0682   0.1788   1.0000
  10.500   1.5213   0.04726   0.03907  -0.0670   0.1762   1.0000
  10.750   1.5367   0.04895   0.04088  -0.0660   0.1741   1.0000
  11.000   1.5555   0.05056   0.04255  -0.0654   0.1722   1.0000
  11.250   1.5621   0.05356   0.04577  -0.0640   0.1706   1.0000
  11.500   1.5390   0.05848   0.05123  -0.0607   0.1693   1.0000
  11.750   1.4975   0.06443   0.05763  -0.0568   0.1686   1.0000
  12.000   1.4182   0.07357   0.06711  -0.0540   0.1688   1.0000
<< Back to WORTMANN FX M2 AIRFOIL (fxm2-il)

Polar data table (+)

Polar graphs


<< Back to WORTMANN FX M2 AIRFOIL (fxm2-il)